Turbofan engine front section

ABSTRACT

A turbofan gas turbine engine includes, among other things, a fan section including a fan hub and an outer housing, the fan hub including a hub diameter supporting a plurality of fan blades, a turbine section including a fan drive turbine, and a geared architecture that interconnects the fan drive turbine and the fan hub, the geared architecture including a gear volume.

CROSS REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.15/007,263, filed on Jan. 27, 2016, which is a continuation of U.S.patent application Ser. No. 14/785,872, filed on Oct. 21, 2015, which isa National Stage Entry of PCT Application No. PCT/US2014/035821, filedon Apr. 29, 2014, which claims priority to U.S. Provisional ApplicationNo. 61/821,371 filed on May 9, 2013.

BACKGROUND

A turbofan engine includes a fan section, a compressor section, acombustor section and a turbine section. Air entering the compressorsection is compressed and delivered into the combustion section where itis mixed with fuel and ignited to generate a high-energy exhaust gasflow. The high-energy exhaust gas flow expands through the turbinesection to drive the compressor and the fan section. The compressorsection typically includes low and high pressure compressors, and theturbine section includes low and high pressure turbines.

A speed reduction device such as an epicyclical gear assembly may beutilized to drive the fan section such that the fan section may rotateat a speed different than the turbine section so as to increase theoverall propulsive efficiency of the engine. The geared architecture maybe located in a front section of the engine and thereby influences howairflow paths are defined to the compressor section. Airflow efficiencyinto the compressor section provides increased overall engine efficiencyand therefore any improvements are desirable.

Turbofan engine manufacturers continue to seek improvements to engineperformance including improvements to thermal, transfer and propulsiveefficiencies.

SUMMARY

A front section for a gas turbine engine according to an example of thepresent disclosure includes a fan section including a fan hub. The fanhub includes a hub diameter supporting a plurality of fan bladesincluding a tip diameter. A ratio of the hub diameter to the tipdiameter is between about 0.24 and about 0.36. A transitional entrancepassage is configured to communicate flow between the fan section and acompressor section. The transitional entrance passage includes an inletdisposed at an inlet diameter and an outlet to the compressor sectiondisposed at an outlet diameter. A ratio of the hub diameter to the inletdiameter is between about 0.65 and about 0.95.

In a further embodiment of any of the foregoing embodiments, theplurality of fan blades is less than twenty (20) fan blades.

In a further embodiment of any of the foregoing embodiments, the ratioof the hub diameter to the inlet diameter is between about 0.70 andabout 0.90.

In a further embodiment of any of the foregoing embodiments, the fansection is configured to deliver air into a bypass duct, and a portionof air into the compressor section, with a bypass ratio defined as thevolume of air delivered into the bypass duct compared to the volume ofair delivered into the compressor section, and the bypass ratio beinggreater than about 10.

In a further embodiment of any of the foregoing embodiments, a pressureratio across the fan section is less than about 1.5.

In a further embodiment of any of the foregoing embodiments, thecompressor section includes a multi-stage low pressure compressor.

A gas turbine engine according to an example of the present disclosureincludes a fan section including a fan hub. The fan hub includes a hubdiameter supporting a plurality of fan blades including a tip diameter,a ratio of the hub diameter to the tip diameter being between about 0.24and about 0.36. A compressor section includes a low pressure compressorand a high pressure compressor. A turbine section includes a fan driveturbine configured to drive the fan section and the low pressurecompressor. The fan drive turbine has a greater number of stages thanthe low pressure compressor. The fan drive turbine has fewer stages thanthe high pressure compressor. A transitional entrance passage couplesthe fan section and the compressor section. The transitional entrancepassage includes an inlet disposed at an inlet diameter and an outletdisposed at an outlet diameter. The inlet is adjacent to the fansection. The outlet is adjacent to the compressor section. A ratio ofthe hub diameter to the inlet diameter is between about 0.65 and about0.95.

In a further embodiment of any of the foregoing embodiments, the lowpressure compressor is a multi-stage compressor.

In a further embodiment of any of the foregoing embodiments, the ratioof the hub diameter to the inlet diameter is between about 0.70 andabout 0.90.

In a further embodiment of any of the foregoing embodiments, the fansection is configured to deliver air into a bypass duct, and a portionof air into the compressor section, with a bypass ratio defined as thevolume of air delivered into the bypass duct compared to the volume ofair delivered into the compressor section. The bypass ratio is greaterthan about 6.

A further embodiment of any of the foregoing embodiments includes ageared architecture configured to rotate the fan hub at a lower relativespeed than the fan drive turbine.

In a further embodiment of any of the foregoing embodiments, theplurality of fan blades is less than twenty (20) fan blades.

In a further embodiment of any of the foregoing embodiments, the lowpressure compressor is a three stage compressor.

In a further embodiment of any of the foregoing embodiments, the ratioof the hub diameter to the inlet diameter is between about 0.70 andabout 0.90.

In a further embodiment of any of the foregoing embodiments, the turbinesection includes a high pressure turbine configured to drive the highpressure compressor, the high pressure turbine including at least twostages.

A method of designing a gas turbine engine according to an example ofthe present disclosure includes providing a fan section including a fanhub. The fan hub includes a hub diameter supporting fewer than 26 fanblades to define a tip diameter. A ratio of the hub diameter to the tipdiameter is between about 0.24 and about 0.36. The method includesproviding a compressor section coupled to a turbine section. The turbinesection is configured to drive the fan section providing a transitionalentrance passage coupling the fan section and the compressor section.The transitional entrance passage includes an inlet disposed at an inletdiameter and an outlet to the compressor section disposed at an outletdiameter. A ratio of the hub diameter to the inlet diameter is betweenabout 0.70 and about 0.90.

In a further embodiment of any of the foregoing embodiments, thecompressor section includes a low pressure compressor and a highpressure compressor. The low pressure compressor is a multi-stagecompressor, and the high pressure compressor has a greater number ofstages than the low pressure compressor.

In a further embodiment of any of the foregoing embodiments, the hubdiameter supports fewer than 20 fan blades.

In a further embodiment of any of the foregoing embodiments, the fansection is configured to deliver air into a bypass duct, and a portionof air into the compressor section, with a bypass ratio defined as thevolume of air delivered into the bypass duct compared to the volume ofair delivered into the compressor section. The bypass ratio is greaterthan about 10.

A further embodiment of any of the foregoing embodiments includesproviding a geared architecture configured to rotate the fan hub at alower relative speed than the turbine section.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example turbofan engine.

FIG. 2 is a schematic view of an example front section of a turbofanengine.

FIG. 3 is a schematic view of another example front section of aturbofan engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example turbofan engine 20 thatincludes a fan section 22 and a core engine section 18 that includes acompressor section 24, a combustor section 26 and a turbine section 28.Alternative engines might include an augmenter section (not shown) amongother systems or features. The fan section 22 drives air along a bypassflow path B while the compressor section 24 draws air in along a coreflow path C where air is compressed and communicated to a combustorsection 26. In the combustor section 26, air is mixed with fuel andignited to generate a high pressure exhaust gas stream that expandsthrough the turbine section 28 where energy is extracted and utilized todrive the fan section 22 and the compressor section 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis and where a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool that enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool that enables ahigh pressure turbine to drive a high pressure compressor of thecompressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about 5. The pressure ratio of the example low pressure turbine 46is measured prior to an inlet of the low pressure turbine 46 as relatedto the pressure measured at the outlet of the low pressure turbine 46prior to an exhaust nozzle.

A mid-turbine frame 58 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the lowpressure turbine 46.

Airflow through the core airflow path C is compressed by the lowpressure compressor 44 then by the high pressure compressor 52 mixedwith fuel and ignited in the combustor 56 to produce high speed exhaustgases that are then expanded through the high pressure turbine 54 andlow pressure turbine 46. The mid-turbine frame 58 includes vanes 60,which are in the core airflow path and function as an inlet guide vanefor the low pressure turbine 46. Utilizing the vane 60 of themid-turbine frame 58 as the inlet guide vane for low pressure turbine 46decreases the length of the low pressure turbine 46 without increasingthe axial length of the mid-turbine frame 58. Reducing or eliminatingthe number of vanes in the low pressure turbine 46 shortens the axiallength of the turbine section 28. Thus, the compactness of the turbofanengine 20 is increased and a higher power density may be achieved.

The disclosed turbofan engine 20 in one example is a high-bypass gearedaircraft engine. In a further example, the turbofan engine 20 includes abypass ratio greater than about six (6), with an example embodimentbeing greater than about ten (10). The example geared architecture 48 isan epicyclical gear train, such as a planetary gear system, star gearsystem or other known gear system, with a gear reduction ratio ofgreater than about 2.3.

In one disclosed embodiment, the turbofan engine 20 includes a bypassratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by airflow through the bypassflow path B due to the high bypass ratio. The fan section 22 of theengine 20 is designed for a particular flight condition—typically cruiseat about 0.8 Mach and about 35,000 feet. The flight condition of 0.8Mach and 35,000 ft., with the engine at its best fuel consumption—alsoknown as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—isthe industry standard parameter of pound-mass (lbm) of fuel per hourbeing burned divided by pound-force (lbf) of thrust the engine producesat that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram°R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed”, as disclosedherein according to one non-limiting embodiment, is less than about 1150ft/second.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example turbofan engine 20with increased power transfer efficiency.

The gas turbofan engine 20 includes a front section 62 extending fromthe fan section 22 axially aft to bearing assembly 108 supporting aforward portion of the low speed spool 30. The front section 62 includesthe fan section 22, the geared architecture 48, and compressor inletpassage 88 part of the core flow path C. The fan section 22 includes afan hub 64 that supports the fan blades 42.

The fan hub 64 supports each of the blades 42 for rotation about theaxis A. Each of the blades 42 includes a blade tip 68. A tip diameter 70is disposed between opposing blade tips 68 and defines the diameter ofthe fan section 22. The blades 42 extend from a root portion 45supported within a fan hub 64. The fan hub 64 defines a hub diameter 66.The hub diameter 66 is related to the tip diameter 70 according to aratio that reflects a size of the bypass flow path B related to the coreengine section 18. In the disclosed embodiment the ratio of hub diameter66 to the tip diameter 70 is between about 0.24 and about 0.36.

The core flow path C includes a compressor inlet passage 88 that isdisposed annularly about the geared architecture 48. The compressorinlet passage 88 includes an inlet 90 into supporting structure for thegeared architecture 48 and the fan hub 64 and an outlet 94 aft of thesupporting structure. The outlet 94 directs air into a first stage ofthe low pressure compressor 44. The hub 64, inlet 90 and outlet 94define a path for air entering the turbofan engine 20 and entering thelow pressure compressor 44.

Referring to FIG. 2, with continued reference to FIG. 1, the inlet 90 isset at an inlet diameter 92 and the outlet 94 is set at an outletdiameter 96. The inlet 90 is at the leading edge and innermost radius ofthe first vane 102 aft of the fan section 22. The outlet 94 is theinnermost radius of the first rotating airfoil 120 of the first or lowpressure compressor 44. The hub diameter 66, inlet diameter 92 andoutlet diameter 96 define a transitional flow path 106 into the lowpressure compressor 44. The transitional flow path 106 includes complexflow fields with swirl components that are turned by vanes 102 and 104within the compressor inlet passage 88 to provide desired axiallyoriented flows.

The transitional flow path 106 includes relatively low diametricalchanges to provide a smooth aerodynamic flow path into the low pressurecompressor 44. The shape of the transitional flow path 106 greatlyimproves and simplifies aerodynamic flow performance through thecompressor inlet passage 88 to increase engine efficiency.

The shape of the transitional flow path 106 is defined by ratios betweenthe hub diameter 66, inlet diameter 92 and outlet diameter 96. Thedisclosed transitional flow path 106 includes a ratio of the inletdiameter 92 to the outlet diameter 96 that is between about 1.10 andabout 1.64, or more narrowly between about 1.24 and about 1.44. Thetransitional flow path 106 further includes a ratio of the hub diameter66 to the inlet diameter 92 that is between about 0.65 and about 0.95,or more narrowly between about 0.70 and about 0.90.

In one example engine embodiment, the hub diameter 66 is between about21.0 inches (53.34 cm) and about 26.0 inches (66.04 cm). The inletdiameter 92 is between about 26.5 inches (67.31 cm) and about 33.0inches (83.82 cm). The outlet diameter 96 is between about 20.0 inches(50.8 cm) and about 24.5 inches (62.23 cm).

The inlet diameter 92 is the largest of the diameters 66, 92 and 96defining the transitional flow path 106 and defines a necessaryinflection point from the convergence of the root portion 45 of the fanblade 42 that provides desired aerodynamic performance.

The transitional flow path 106 between the hub diameter 66, the inletdiameter 92 and outlet diameter 96 is enabled by a gear diameter 98defined by the geared architecture 48 and by the axial and radialposition of the forward bearing assembly 108. The forward bearingassembly 108 is positioned axially and radially relative to the lowpressure compressor 44 to enable the subtle changes in the transitionalflow path 106. Accordingly, the inlet diameter 92, and therefore thedesired inflection point is enabled by the size of the gearedarchitecture 48, and the outlet diameter 96 is enabled by the size andposition of the forward bearing assembly 108.

The geared architecture 48 includes a sun gear 76 driven by the lowpressure turbine 46 through the inner shaft 40. The sun gear 76 drivesintermediate planetary gears (either planet gears or star gears) 78supported for rotation by journal bearings 100. A carrier 82 supportsthe journal bearings 100 and thereby the planetary gears 78. A ring gear80 circumscribes the planetary gears 78. In this example, the ring gear80 is attached to drive the fan hub 64 about the axis A. The carrier 82is fixed and holds the intermediate planetary gears 78 from rotationabout the axis A.

The geared architecture 48 illustrated in FIG. 1 is a star epicyclicalconfiguration where the ring gear 80 is driven about the axis A and thecarrier 80 is fixed to a portion of the engine static structure 36.However, other gear configurations are within the contemplation of thisdisclosure.

Referring to FIG. 3, another geared architecture 49 is shown thatincludes a sun gear 75 that drives planet gears 79 supported in acarrier 83 that is attached to drive a fan hub 65. A ring gear 81circumscribes the planet gears 79 and is fixed to a portion of theengine static structure 36. The geared architecture 49 drives the fanhub 65 through rotation of the planet gears 79 and carrier 83 about theaxis A and is referred to as a planet epicyclical gear configuration.The disclosed features and size are applicable to either of thedisclosed geared architectures 48, 49 illustrated schematically in FIGS.2 and 3. Further explanation and disclosure are explained with regard tothe geared architecture 48 illustrated in FIG. 2, but is just asapplicable to the embodiment illustrated and explained in FIG. 3.

Referring back to FIGS. 1 and 2, the carrier 82 includes an outerperiphery 85 and the ring gear 80 includes the gear diameter 98 thatcombines to define a gear volume 72. The gear diameter 98 defined by thering gear 80 and carrier 82 define the boundary of the gear volume 72.The gear volume 72 includes the elements of the geared architecture suchas the journal bearings, 100, carrier 80, sun gear 76, planetary gears78 and ring gear 80. The gear volume 72 does not encompass the mountingand flexible structures that may be utilized to support the gearedarchitecture.

The gear volume 72 is the annular volume about the axis A, definedwithin the bounds of the gear diameter 98 and axial length 74. The axiallength 74 of the geared architecture 48 includes the carrier 82. In thedisclosed example, the geared architecture 48 includes an axial length74 between about 4.16 inches (10.56 cm) and about 6.90 inches (17.53cm).

The geared architecture 48 provides for the transmission of inputhorsepower 84 from the low pressure turbine 46 to output horsepower 86to the fan section 22. The efficient transmission of input horsepower 84is enabled by the configuration defining the gear volume 72. In thisexample, the gear volume is between about 1318 in³ (21,598 cm³) andabout 1977 in³ (32,397 cm³).

The gearbox volume is necessary for the transfer of power from the fandrive or low pressure turbine 46 to the fan section 22. The geardiameter 98 is held close to the fan hub diameter 66 to define theflowpath 106 to be as short and unvarying in diameter as possible. Theshort and unvarying diameter of the transitional flow path 106 enablespreferred pressure recovery between the fan blade root 45, the inlet 92and the outlet 94 to the first rotating stage of the first or lowpressure compressor 44.

In one example, the range of gear volume 72 is provided for an engine 20that generates thrust ranging between about 21,000 lbf (93,412 N) and35,300 lbf (157,022 N). The thrust generated is a function of theefficiency of the engine configuration and of the transfer of horsepowerthrough the geared architecture 48. A measure of the efficiency of thegeared architecture for a give volume is a power transfer parameter(PTP) and is defined as the power transferred through the gearedarchitecture 48 divided by the gear volume 72 and multiplied by anoverall gear ratio, a set out in Equation 1.

Power Transfer Parameter=[Power Transferred (HP)/Gear Volume(in³)]×overall gear ratio.  Equation 1:

The PTP provides a normalized factor for comparison between gearedarchitectures for different engine configurations. Moreover the gearratio accounts for the extra work performed for higher gear ratios.Embodiments of the geared turbofan engine that include the disclosedgeared architecture 48 gear volumes 72 and that generate thrust rangingbetween about 21,000 lbf (93,412 N) and 35,300 lbf (157,022 N) include aPTP of between about 430 and about 645.

The PTP of the example geared architecture 48 enables increased transferof power while maintaining a size and volume that further enables thetransitional flow path 106 orientations that provide desired aerodynamicflow properties.

The forward bearing assembly 108 is disposed at an axial distance 110from the outlet 94 to support rotation of the forward portion of the lowspeed spool 30 including the low pressure compressor 44. The position ofthe forward bearing assembly 108 provides a desired balance androtational support of the low pressure compressor 44. Placement of theforward bearing assembly 108 is desired within a mid-region of thecompressor 44 and requires a radial space sufficient to supportlubricant and cooling features required during operation. Accordingly, adiameter 112 of the bearing assembly 108 has a direct effect on theconfiguration of the low pressure compressor 44 and thereby the positionof the outlet 94. Moreover, the axial distance 118 from the forward tipof the hub 64 to the bearing assembly 108 is enabled by the size andvolume of the geared architecture 48 and combined with the position ofthe forward bearing assembly 108 enables the desirable design of thetransitional flow path 106.

In one disclosed dimensional engine embodiment the diameter 112 measuredto a center point of the bearing assembly 108 is between about 7.25inches (18.4 cm) and about 9.0 inches (22.86 cm). The axial distance 110is between about 4.25 inches (10.79 cm) and 9.00 inches (22.86 cm). Anoverall axial length 118 of the front section 62 from the hub diameter66 at the forward portion of the fan hub 64 to the forward bearingassembly is between about 33.00 inches (83.82 cm) and about 67.50 inches(171.45 cm). The axial distance 110 between the outlet 94 and thebearing assembly 108 enable the desired reduced length of the forwardsection 62.

The disclosed dimensional embodiment is only one example enabling thedisclosed configuration of the transitional flow path 106. Theconfiguration of disclosed engine embodiments is reflected in a ratio ofthe overall length 118 to the axial distance 110 that is between about3.6 and 15.8. Moreover, a ratio between the outlet diameter 96 and thebearing assembly diameter 112 is between about 2.25 and 3.35. Theseratios reflect the configuration that enables the radial and axialposition of the outlet 94.

The axial length 74 of the geared architecture 48 further enables thedesired relatively flat transitional flow path 106. The volume of thegeared architecture 48 enables the power transfer to the fan hub 64 andis a factor determined by the axial length 74 and the gear diameter 98.Decreasing the gear diameter 98 enables a corresponding reduction inaxial length 74 that in turn enables the desired configuration of thetransitional flow path 106.

Therefore, a relationship between the axial length 74 of the gearedarchitecture and the overall length 118 of the front section 62 furtherreflects the disclosed configuration of the transitional flow path 106and engine 20. A ratio of the overall length 118 as related to the axiallength 74 of the geared architecture 48 is between about 4.8 and about16.2. The ratio of the overall length 118 to the axial length 74reflects the disclosed geared architecture 48 including the geardiameter 98 and volume 72 that the desired configuration of thetransitional flow path 106 and front section 62.

Accordingly, the gear volume 72, gear diameter 98, and axial length 74of the geared architecture along with the location of the forwardbearing assembly 108 enable an efficient transitional flow path 106 inthe disclosed compact front section 62.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. A turbofan gas turbine engine comprising: a fansection including a fan hub and an outer housing, the fan hub includinga hub diameter supporting a plurality of fan blades, the outer housingsurrounding the fan blades to establish a bypass duct, the fan bladesincluding a tip diameter with a ratio of the hub diameter to the tipdiameter between 0.24 and 0.36, and a pressure ratio across the fansection of less than 1.45 across the fan blades alone at cruise at 0.8Mach and 35,000 feet; a compressor section including a first compressorand a second compressor; a turbine section including a fan drive turbineand a second turbine, wherein the second turbine drives the secondcompressor; a geared architecture that interconnects the fan driveturbine and the fan hub, the geared architecture including a gear volumebetween 1318 in³ and 1977 in³; and a compressor inlet passage disposedannularly about the geared architecture, the compressor inlet passagecoupling the fan section and the compressor section, the compressorinlet passage including an inlet disposed at an inlet diameter and anoutlet disposed at an outlet diameter, the inlet adjacent to the fansection, the outlet adjacent to the compressor section, and a ratio ofthe hub diameter to the inlet diameter being between 0.65 and 0.95. 2.The turbofan gas turbine engine as recited in claim 1, wherein thegeared architecture comprises an epicyclic gear system including a sungear, a ring gear, a plurality of intermediate gears and a carrier,wherein the ring gear circumscribes the intermediate gears, theintermediate gears are driven by the sun gear, and the carrier supportsthe intermediate gears, and the gear volume is defined within a spacebounded by the ring gear and an outer periphery of the carrier.
 3. Theturbofan gas turbine engine as recited in claim 2, further comprising abypass ratio of greater than
 10. 4. The turbofan gas turbine engine asrecited in claim 3, wherein the second turbine includes 2 stages.
 5. Theturbofan gas turbine engine as recited in claim 4, wherein both thefirst compressor and the fan drive turbine include a greater number ofstages than the second turbine, and the second compressor includes agreater number of stages than the first compressor.
 6. The turbofan gasturbine engine as recited in claim 5, wherein the turbofan gas turbineengine is sized to generate thrust ranging between 21,000 lbf and 35,300lbf.
 7. The turbofan gas turbine engine as recited in claim 6, wherein apower transfer parameter is defined as power transferred through thegeared architecture divided by the gear volume multiplied by a gearreduction ratio, and the power transfer parameter is between 430 and645.
 8. The turbofan gas turbine engine as recited in claim 7, whereinthe gear reduction ratio is greater than 2.3.
 9. The turbofan gasturbine engine as recited in claim 8, wherein the fan has less than 20fan blades.
 10. The turbofan gas turbine engine as recited in claim 9,wherein the number of turbine rotors of the fan drive turbine is between3.3 and 8.6.
 11. The turbofan gas turbine engine as recited in claim 10,wherein the fan drive turbine includes 5 stages.
 12. The turbofan gasturbine engine as recited in claim 10, wherein the inlet diameter isestablished at a leading edge and innermost radius of a first vane aftof the fan section, the outlet diameter is established at an innermostradius of a forwardmost rotating airfoil of the first compressor, and aratio of the inlet diameter to the outlet diameter is between 1.10 and1.64.
 13. The turbofan gas turbine engine as recited in claim 12,wherein the ratio of the hub diameter to the inlet diameter is between0.70 and 0.90.
 14. The turbofan gas turbine engine as recited in claim13, wherein the geared architecture includes an outer diameter that isless than the inlet diameter of the compressor inlet passage.
 15. Theturbofan gas turbine engine as recited in claim 9, wherein the gearedarchitecture is a star gear system, the ring gear drives the fan hub,and the carrier is fixed to an engine static structure.
 16. The turbofangas turbine engine as recited in claim 15, wherein the inlet diameter isestablished at a leading edge and innermost radius of a first vane aftof the fan section, the outlet diameter is established at an innermostradius of a forwardmost rotating airfoil of the first compressor, and aratio of the inlet diameter to the outlet diameter is between 1.10 and1.64.
 17. The turbofan gas turbine engine as recited in claim 16,further comprising: a forward bearing assembly that supports a forwardportion of a first shaft that drives the first compressor; wherein theturbofan gas turbine engine includes an overall axial distance from aforward part of the fan hub to the forward bearing assembly, and a ratioof the overall axial distance to an axial length of the gearedarchitecture is between 4.8 and 16.2.
 18. The turbofan gas turbineengine as recited in claim 17, further comprising: a mid-turbine framebetween the fan drive turbine and the second turbine, the mid-turbineframe supporting bearing systems in the turbine section and includingvanes in a core flow path; and wherein the fan drive turbine drives boththe sun gear and the first compressor.
 19. The turbofan gas turbineengine as recited in claim 18, wherein: the ratio of the hub diameter tothe inlet diameter is between 0.70 and 0.90; a ratio between the overallaxial distance and an axial length between the forward bearing assemblyand the outlet to the compressor inlet passage is between 3.6 and 15.8;the forward bearing assembly is disposed at a diameter, and a ratio ofthe outlet diameter to the diameter of the forward bearing assembly isbetween 2.25 and 3.35; the ratio of the inlet diameter to the outletdiameter is between 1.24 and 1.44; and the axial length of the gearedarchitecture is between 4.16 inches and 6.90 inches.
 20. The turbofangas turbine engine as recited in claim 9, wherein the gearedarchitecture is a planetary gear system, the carrier drives the fan hub,and the ring gear is fixed to an engine static structure.
 21. Theturbofan gas turbine engine as recited in claim 20, wherein a ratiobetween the number of the fan blades and the number of turbine rotors ofthe fan drive turbine is between 3.3 and 8.6.
 22. The turbofan gasturbine engine as recited in claim 21, wherein the inlet diameter isestablished at a leading edge and innermost radius of a first vane aftof the fan section, the outlet diameter is established at an innermostradius of a forwardmost rotating airfoil of the first compressor, and aratio of the inlet diameter to the outlet diameter is between 1.10 and1.64.
 23. The turbofan gas turbine engine as recited in claim 22,further comprising: a forward bearing assembly that supports a forwardportion of a first shaft that drives the first compressor; wherein theturbofan gas turbine engine includes an overall axial distance from aforward part of the fan hub to the forward bearing assembly, and a ratioof the overall axial distance to an axial length of the gearedarchitecture is between 4.8 and 16.2.
 24. The turbofan gas turbineengine as recited in claim 23, wherein the axial length of the gearedarchitecture is between 4.16 inches and 6.90 inches.
 25. The turbofangas turbine engine as recited in claim 23, wherein the ratio of the hubdiameter to the inlet diameter is between 0.70 and 0.90.
 26. Theturbofan gas turbine engine as recited in claim 23, wherein a ratiobetween the overall axial distance and an axial length between theforward bearing assembly and the outlet to the compressor inlet passageis between 3.6 and 15.8.
 27. The turbofan gas turbine engine as recitedin claim 23, wherein the forward bearing assembly is disposed at adiameter, and a ratio of the outlet diameter to the diameter of theforward bearing assembly is between 2.25 and 3.35.
 28. The turbofan gasturbine engine as recited in claim 23, wherein the ratio of the inletdiameter to the outlet diameter is between 1.24 and 1.44.
 29. Theturbofan gas turbine engine as recited in claim 23, further comprising:a mid-turbine frame between the fan drive turbine and the secondturbine, the mid-turbine frame supporting bearing systems in the turbinesection and including vanes in a core flow path; and wherein the fandrive turbine drives both the sun gear and the first compressor.
 30. Theturbofan gas turbine engine as recited in claim 29, wherein: the ratioof the hub diameter to the inlet diameter is between 0.70 and 0.90; aratio between the overall axial distance and an axial length between theforward bearing assembly and the outlet to the compressor inlet passageis between 3.6 and 15.8; the forward bearing assembly is disposed at adiameter, and a ratio of the outlet diameter to the diameter of theforward bearing assembly is between 2.25 and 3.35; the ratio of theinlet diameter to the outlet diameter is between 1.24 and 1.44; and theaxial length of the geared architecture is between 4.16 inches and 6.90inches.